Fairing Assembly

ABSTRACT

Fairing assemblies and methods for assembling gas turbine engine fairing assemblies are provided. For example, a fairing assembly comprises a plurality of fairings, an annular inner band defining a plurality of inner pockets, and an annular outer band defining a plurality of outer pockets. Each fairing has an inner end radially spaced apart from an outer end. Each inner pocket is shaped complementary to each fairing inner end and has forward and aft ends. Each outer pocket is shaped complementary to each fairing outer end and has forward and aft ends. The inner and outer bands are each a single piece structure. Each fairing inner end is received within one of the plurality of inner pockets, and each fairing outer end is received within one of the plurality of outer pockets. Some embodiments also comprise an inner ring positioned against the inner band to close the inner pockets.

FIELD

The present subject matter relates generally to gas turbine engines.More particularly, the present subject matter relates to airfoilassemblies for gas turbine engines, such as fairing assemblies. Mostparticularly, the present subject matter relates to composite fairingassemblies.

BACKGROUND

More commonly, non-traditional high temperature composite materials,such as ceramic matrix composite (CMC) materials, are being used inapplications such as gas turbine engines. Components fabricated fromsuch materials have a higher temperature capability compared withtypical components, e.g., metal components, which may allow improvedcomponent performance and/or increased engine temperatures. Compositecomponents may provide other advantages as well, such as an improvedstrength to weight ratio.

Typically, a CMC turbine nozzle fairing comprises an airfoil, an innerband, and an outer band that are integrally formed as a single componentthat is split axially into forward and aft sections and splitcircumferentially into a plurality of segments. The plurality ofsegments together form an annular fairing assembly and splitting thecomponent into forward and aft sections allows the airfoil portion to beinstalled around structural elements of the turbine frame, such asstruts or the like. Although splitting the fairing assembly intosections and segments enables assembly with the turbine frame, havingforward and aft sections for each of the plurality of segments increasesa part count for the fairing assembly. Further, splitting the componentincreases the likelihood for leakages, e.g., between eachcircumferential fairing segment, which may also increase the part countby requiring seals or other mechanisms in an attempt to prevent suchleakage. Additionally, thermal differences, i.e., a thermal fight,between the airfoil and bands produce high stresses in the nozzlefairings, which limits the acceptability of part defects and results intighter inspection limits for non-destructive examination of the parts.

Accordingly, improved fairing assemblies would be useful. In particular,a fairing assembly comprising a plurality of fairing airfoils that areeach separate from each of an annular, single piece inner band and anannular, single piece outer band would be advantageous. Further, afairing assembly having single piece annular inner and outer bands thatare adapted for use with multiple turbine frame configurations would bedesirable.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present subject matter, a fairingassembly for a gas turbine engine is provided. The fairing assemblycomprises a plurality of fairings, an annular inner band defining aplurality of inner pockets, and an annular outer band defining aplurality of outer pockets. Each fairing has an inner end radiallyspaced apart from an outer end and extends axially from a leading edgeto a trailing edge. Each inner pocket is shaped complementary to theinner end of each fairing and has a forward end and an aft end. Eachouter pocket is shaped complementary to the outer end of each fairingand has a forward end and an aft end. The inner band is a single piecestructure, and the outer band is a single piece structure. The inner endof each fairing is received with an inner pocket of the plurality ofinner pockets, and the outer end of each fairing is received within anouter pocket of the plurality of outer pockets.

In another exemplary embodiment of the present subject matter, a fairingassembly for a gas turbine engine is provided. The fairing assemblycomprises a plurality of fairings, an inner ring defining a plurality ofinner pocket forward segments, an annular inner band defining aplurality of inner pocket aft segments, and an annular outer banddefining a plurality of outer pockets. Each fairing has an inner endradially spaced apart from an outer end and extends axially from aleading edge to a trailing edge. The inner ring is positioned against aforward edge of the inner band such that the inner pocket forwardsegments and inner pocket aft segments form a plurality of innerpockets. Each inner pocket is shaped complementary to the inner end ofeach fairing and has a forward end and an aft end. Further, each outerpocket is shaped complementary to the outer end of each fairing and hasa forward end and an aft end. The inner ring is a single piecestructure, the inner band is a single piece structure, and the outerband is a single piece structure. The inner end of each fairing isreceived with an inner pocket of the plurality of inner pockets, and theouter end of each fairing is received within an outer pocket of theplurality of outer pockets.

In a further exemplary embodiment of the present subject matter, amethod for assembling a fairing assembly in a gas turbine engine isprovided. The method comprises installing an annular inner band in thegas turbine engine. The inner band defines a plurality of inner pockets.The method further comprises inserting an inner end of each of aplurality of fairings into an inner pocket of the plurality of innerpockets. The method also comprises sliding an annular outer band intoposition with respect to the plurality of fairings such than an outerend of each of the plurality of fairings is received in an outer pocketof a plurality of outer pockets defined by the outer band. The innerband is a single piece structure, and the outer band is a single piecestructure.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-section view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIGS. 2A and 2B provide schematic cross-section views of a fairingassembly according to exemplary embodiments of the present subjectmatter.

FIG. 3 provides a perspective view of the fairing assembly of FIG. 2A orFIG. 2B.

FIG. 4 provides an exploded perspective view of an inner band and anouter band of the fairing assembly of FIG. 2A or FIG. 2B.

FIG. 5 provides a perspective view of a fairing airfoil of the fairingassembly of FIG. 2A or FIG. 2B.

FIGS. 6A and 6B provide schematic cross-section views of a fairingassembly according to exemplary embodiments of the present subjectmatter.

FIG. 7 provides a perspective view of the fairing assembly of FIG. 6A.

FIG. 8 provides an exploded perspective view of an inner band and aninner ring of the fairing assembly of FIG. 6A.

FIG. 9 provides a schematic cross-section view of a fairing assemblyaccording to an exemplary embodiment of the present subject matter.

FIG. 10 provides a perspective view of the fairing assembly of FIG. 9.

FIG. 11 provides an exploded perspective view of an outer band and anouter ring of the fairing assembly of FIG. 9.

FIG. 12 provides an exploded perspective view of an axially splitfairing airfoil according to an exemplary embodiment of the presentsubject matter.

FIG. 13 provides a schematic cross-section view of a circumferentiallysplit fairing airfoil according to an exemplary embodiment of thepresent subject matter.

FIG. 14 provides a perspective view of a portion of a fairing assemblyaccording to an exemplary embodiment of the present subject matter.

FIGS. 15, 16, and 17 provide radial cross-section views of a portion ofa fairing outer end and an outer band pocket with a seal positionedtherebetween, according to various exemplary embodiments of the presentsubject matter.

FIG. 18A provides a perspective view of a pin for pinning a fairingairfoil within a fairing assembly.

FIG. 18B provides an axial cross-section view of a split fairing airfoilhaving two pins as shown in FIG. 18A received therein, according to anexemplary embodiment of the present subject matter.

FIG. 19 provides a schematic cross-section view of a fairing assemblyhaving pinned fairing airfoils, according to an exemplary embodiment ofthe present subject matter.

FIGS. 20A, 20B, and 21-24 provide schematic cross-section views ofgrommet and fastener configurations, according to various exemplaryembodiments of the present subject matter.

FIGS. 25, 26, and 27 provide flow diagrams illustrating methods forassembling a fairing assembly in a gas turbine engine, according tovarious exemplary embodiments of the present subject matter.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

As used herein, the terms “first,” “second,” and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a,” “an,” and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about,” “approximately,” and “substantially,” are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 10percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

For the depicted embodiment, fan section 14 includes a fan 38 having aplurality of fan blades 40 coupled to a disk 42 in a spaced apartmanner. As depicted, fan blades 40 extend outward from disk 42 generallyalong the radial direction R. The fan blades 40 and disk 42 are togetherrotatable about the longitudinal axis 12 by LP shaft 36. In someembodiments, a power gear box having a plurality of gears may beincluded for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

In some embodiments, components of turbofan engine 10, particularlycomponents within or defining the hot gas path 78, may comprise acomposite material, such as a ceramic matrix composite (CMC) materialhaving high temperature capability. Composite materials generallycomprise a fibrous reinforcement material embedded in matrix material,e.g., a ceramic matrix material. The reinforcement material serves as aload-bearing constituent of the composite material, while the matrix ofa composite material serves to bind the fibers together and act as themedium by which an externally applied stress is transmitted anddistributed to the fibers.

Exemplary CMC materials may include silicon carbide (SiC), silicon,silica, or alumina matrix materials and combinations thereof. Ceramicfibers may be embedded within the matrix, such as oxidation stablereinforcing fibers including monofilaments like sapphire and siliconcarbide (e.g., Textron's SCS-6), as well as rovings and yarn includingsilicon carbide (e.g., Nippon Carbon's NICALON®, Ube Industries'TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates (e.g.,Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's440 and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si,Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g.,pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite).For example, in certain embodiments, bundles of the fibers, which mayinclude a ceramic refractory material coating, are formed as areinforced tape, such as a unidirectional reinforced tape. A pluralityof the tapes may be laid up together (e.g., as plies) to form a preformcomponent. The bundles of fibers may be impregnated with a slurrycomposition prior to forming the preform or after formation of thepreform. The preform may then undergo thermal processing, such as a cureor burn-out to yield a high char residue in the preform, and subsequentchemical processing, such as melt-infiltration with silicon, to arriveat a component formed of a CMC material having a desired chemicalcomposition. In other embodiments, the CMC material may be formed as,e.g., a carbon fiber cloth rather than as a tape.

As stated, components comprising a composite material may be used withinthe hot gas path 78, such as within the combustion and/or turbinesections of engine 10. As an example, one or more stages of turbinerotor blades and/or turbine nozzles may be CMC components formed fromCMC materials. However, composite components made from CMC or othercomposite materials may be used in other sections as well, e.g., thecompressor and/or fan sections.

Turning to FIGS. 2A and 2B, schematic cross-section views are providedof a fairing assembly 100 according to exemplary embodiments of thepresent subject matter. FIG. 3 provides a perspective view of thefairing assembly 100, FIG. 4 provides an exploded perspective view of aninner band 102 and outer band 104 of the fairing assembly 100, and FIG.5 provides a perspective view of a fairing 106 of the fairing assembly100. The fairing assembly 100 includes the inner band 102, the outerband 104 encircling the inner band 102, and a plurality of vanes orfairings 106 extending between the inner band 102 and the outer band104. At least a portion of the fairings 106 are hollow and encircle orencase a plurality of struts 108, which are part of a turbine frame thatincludes an inner support structure 140, such as an inner hub, and anouter support structure 146, such as an outer case. As shown in FIGS. 2Aand 2B, the inner and outer bands 102, 104 also define openingstherethrough such that the struts 108 extend from the inner supportstructure 140 to the outer support structure 146. Conduits may runthrough some of the struts 108, and additional structures such ashangers and retainers may be included in the fairing assembly 100, e.g.,for attaching the fairing assembly 100 to the engine casing, etc.

The fairing assembly 100 crosses the combustion gas flow path 78 and,thus, in operation is exposed to high temperatures. That is, in thedepicted embodiment, the fairing assembly 100 is a turbine nozzlefairing assembly forming an annular turbine nozzle stage, e.g., aplurality of turbine nozzles positioned circumferentially about theaxial centerline 12 of the engine 10. As such, each of the plurality offairings 106, inner band 102, and outer band 104 form a liner along thehot gas path 78, protecting metallic components and the like from theheat of the combustion gases 66. Further, the exemplary fairing assembly100 described herein forms a transition from the high pressure turbine28 to the low pressure turbine 30 and, as such, each of the inner band102 and outer band 104 are generally conical in shape, increasing incircumference from a forward end 110 of the fairing assembly 100 to anaft end 112 of the assembly 100.

As further shown in FIGS. 2A-5, the inner and outer bands 102, 104 areannular, single piece structures that are separate from each fairing106. That is, each of the inner band 102, outer band 104, and pluralityof fairings 106 are separately formed such that each component is anindividual piece. In exemplary embodiments, the plurality of fairings106, the inner band 102, and the outer band 104 are each formed from acomposite material, such as a CMC material. However, in otherembodiments, the inner band 102, outer band 104, and fairings 106 may bemanufactured from any suitable material using any suitable process ortechnique. For instance, one or more of the components 102, 104, 106 maybe formed from plies of a composite material in a process that includeslaying up, debulking, curing, and densifying the composite; one or moreof the components 102, 104, 106 may be formed from a suitable materialin an additive manufacturing process; and/or one or more of thecomponents 102, 104, 106 may be formed from a metallic material using acasting process or the like.

Keeping with FIGS. 2A-5, there may be any number of vanes or fairings106 included in the fairing assembly 100. The fairings 106 may haveairfoil shapes and may create an airfoil cascade. More particularly,each fairing airfoil 106 may have a concave pressure side 114 opposite aconvex suction side 116, and each side 114, 116 may extend axially froma leading edge 118 to a trailing edge 120. Further, each fairing 106 mayhave an inner end 122 that is radially spaced apart from an outer end124. During operation, the fairings 106 shape the air flow to improvethe engine efficiency. The struts 108, which typically are not anairfoil shape, would negatively impact the airflow and usually areconstructed of a material not capable of withstanding flow pathconditions; therefore, the fairings 106 are included to form an airfoilaround the struts 108. It will be understood that in the illustratedexample, a portion of the fairings 106 surround structural elements,like the struts 108, while the remaining fairings 106 surround nothingstructural. However, as shown in the figures, each fairing 106 may behollow, defining a cavity 202 such that each fairing 106 has an internalcavity pressure as described in greater detail herein.

As shown in FIG. 4, the annular inner band 102 defines a plurality ofinner pockets 126, and the annular outer band 104 defines a plurality ofouter pockets 128. Each inner pocket 126 is shaped complementary to theinner end 122 of each fairing 106, and each outer pocket 128 is shapedcomplementary to the outer end 124 of each fairing 106. Further, eachinner pocket 126 has a forward end 130 and an aft end 132, and eachouter pocket 128 has a forward end 134 and an aft end 136. Asillustrated in FIGS. 2A, 2B, and 3, the inner end 122 of each fairing106 is received with an inner pocket 126 of the plurality of innerpockets 126, and the outer end 124 of each fairing 106 is receivedwithin an outer pocket 128 of the plurality of outer pockets 128. Itwill be appreciated that the fairing assembly 100 comprises an equalnumber of inner pockets 126, outer pockets 128, and fairings 106, withone fairing 106 positioned within one inner pocket 126 and acorresponding outer pocket 128. Moreover, for each fairing 106, theleading edge 118 of the fairing 106 is positioned at the forward end 130of the inner pocket 126 and the forward end 134 of the outer pocket 128in which the fairing 106 is received, and the trailing edge 120 of thefairing 106 is positioned at the aft end 132 of the inner pocket 126 andthe aft end 136 of the outer pocket 128. Additionally, as depicted inFIG. 5, the inner end 122 of each fairing 106 may define an inner boss123 and the outer end 124 of each fairing 106 may define an outer boss125. The inner and outer pockets 126, 128 may be shaped to complementthe inner and outer bosses 123, 125, respectively, such that the bosses123, 125 are received in the pockets 126, 128. Further, each boss 123,125 may provide an area for receipt of a seal, e.g., inner boss 123 mayreceive inner seal 278 and outer boss 125 may receive outer seal 280 asdescribed in greater detail herein.

In the exemplary embodiment illustrated in FIGS. 2A-4, the inner band102 surrounds each inner pocket 126 such that each inner pocket 126 isclosed at its forward end 130 and aft end 132. That is, the innerpockets 126 are not open at either the forward end 110 or aft end 112 ofthe fairing assembly 100. Similarly, the outer band 104 surrounds eachouter pocket 128 such that each outer pocket 128 is closed at itsforward end 134 and aft end 136 and is not open at either the forwardend 110 or aft end 112 of the fairing assembly 100. Accordingly, theillustrated fairing assembly 100 is compatible with a bolted frame(separable hub, strut, and case) or similar frame designs. Moreparticularly, the fairing assembly 100 shown in FIGS. 2A-4 is suitablefor a turbine frame that allows the single piece annular inner band 102to be installed with respect to the frame, the fairings 106 positionedin the inner pockets 126, and the single piece annular outer band 104 tobe slid into position such that the outer ends 124 of the fairings 106are received in the outer pockets 128. Methods of assembling the fairingassembly 100 will be described in greater detail below.

Referring to FIG. 2A, in some embodiments the fairing airfoils 106 maybe brazed to the inner band 102 and/or outer band 104. In an exemplaryembodiment, each fairing 106 is brazed to the inner band 102 at or nearthe inner end 122 of the fairing 106 and each fairing 106 is brazed tothe outer band 104 at or near the outer end 124 of the fairing 106. Inother embodiments, the fairings 106 may be brazed at only one of theinner end 122 or outer end 124. It will be appreciated that, by brazingthe fairing airfoils 106 to the inner and/or outer bands 102, 104, theneed for seals between the bands 102, 104 and fairings 106 iseliminated. Further, brazing is not limited to embodiments in which thebands 102, 104 and fairings 106 are metallic, but also, e.g., brazingmay be used where the bands 102, 104 and fairings 106 are CMC.

As further illustrated in FIGS. 2B and 3, each of the inner band 102 andouter band 104 may be pinned at the forward end 110 of the fairingassembly 100 to retain the assembly 100 in its position within theengine 10. More specifically, a plurality of radially extending innerpins 138 extend from a support structure 140 into a forward portion 142of the inner band 102. Similarly, a plurality of radially extendingouter pins 144 extend from a support structure 146 into a forwardportion 148 of the outer band 104. The support structures 140, 146 maybe hangers, retainers, or the like. In some embodiments, as previouslydescribed, the inner support structure 140 may be the hub of the turbineframe, and the outer support structure 146 may be the case of theturbine frame. As shown in FIGS. 2B and 3, the inner pins 138 may bereceived within openings 150 in the inner band 102, and the outer pins144 may be received within openings 152 in the outer band 104. In someembodiments, the pins 138, 144 may be secured in the openings 150, 152using grommets or the like such that the pins 138, 144 do not extendthrough the respective band 102, 104. In other embodiments, the pins138, 144 may extend through the forward portion 142, 148 of therespective band 102, 104. Using radial pins 138, 144 allow for radialthermal expansion, e.g., of metallic support structures 140, 146, whileconstraining the fairing assembly 100 axially and tangentially. Such aretention configuration alleviates stresses and strains that wouldotherwise arise from a thermal mismatch between the fairing assembly 100and the support structures 140, 146, e.g., where the fairing assembly100 is formed from a composite material and the support structures 140,146 are formed from metals or metal alloys, while still holding thefairing assembly 100 in position within the engine 10.

However, in some embodiments, the fairing assembly 100 may be bolted, orotherwise fastened, rather than pinned at its forward end 110. Inparticular embodiments, the assembly 100 may be bolted to one or moreflexible metal hangers, which flex to compensate for the thermal growthof the metal hanger(s) relative to the assembly 100, which may be formedfrom a composite material and have a different rate of thermalexpansion. In further embodiments, the fairing airfoils 106 may includea feature that allows the fairings 106 to be bolted to a hanger. Othermeans for retaining the fairing assembly 100 in its position within theengine 10 may be used as well.

Further, as shown in FIG. 2A, in embodiments in which the fairingairfoils 106 are brazed to the inner band 102 and outer band 104, theradial inner pins 138 or the radial outer pins 144 may be eliminated.More particularly, by brazing the fairings 106 to the bands 102, 104,the fairing assembly 100 is effectively a single piece structure, andthus, the assembly 100 need only be pinned at the inner band 102 orouter band 104. Although shown in FIG. 2A with outer pins 144, it willbe understood that, in alternative embodiments, the outer pins 144 maybe eliminated and the inner pins 138 may be used to pin the assembly 100at its forward end 110. Additionally, the inner pins 138 or outer pins144 may be eliminated in embodiments in which the fairing airfoils 106are otherwise attached at the inner band 102 and outer band 104, e.g.,where the fairings 106 are pinned to the bands 102, 104 as describedherein.

Turning now to FIGS. 6A-8, other exemplary embodiments of the fairingassembly 100 will be described. FIGS. 6A and 6B each provide a schematiccross-section view of a fairing assembly 100, and FIG. 7 provides aperspective view of the fairing assembly 100 of FIG. 6A. FIG. 8 providesan exploded perspective view of the inner band 102 and an inner ring 154of the fairing assembly 100 of FIG. 6A. Although not depicted in FIGS.6A and 6B, it will be appreciated that the fairing assembly 100 will beinstalled with a turbine frame like the turbine frame shown anddescribed with respect to FIGS. 2A and 2B, although the turbine frame ofFIGS. 6A and 6B may be configured differently from the turbine framedepicted in FIGS. 2A and 2B. For example, the fairing assembly 100depicted in FIGS. 6A-8 may be used with a two-piece turbine frame, orsimilar frame design, rather than a bolted turbine frame or the like asdescribed above with respect to FIGS. 2A-5. An exemplary two-piece framecomprises a hub and strut assembly and a separable outer case, but othertwo-piece frames also may be used.

As depicted in FIGS. 6A-8, the fairing assembly 100 may comprise anannular, single piece inner ring 154 in addition to the annular, singlepiece inner band 102. More particularly, in some embodiments each innerpocket 126 is open at its forward end 130 and closed at its aft end 132.As such, the inner band 102 defines a plurality of inner pocket aftsegments 156. The inner ring 154 is positioned against a forward edge160 of the inner band 102 to close the forward end 130 of each innerpocket 126. Further, the inner ring 154 defines a plurality of innerpocket forward segments 158, and each inner pocket forward segment 158in the inner ring 154 has a corresponding inner pocket aft segment 156in the inner band 102. The inner pocket forward segments 158 align withthe inner pocket aft segments 156 such that each inner pocket forwardsegment 158 and its corresponding inner pocket aft segment 156 togetherform one of the plurality of inner pockets 126. Thus, the inner pockets126 are split, with a forward portion (i.e., inner pocket forwardsegment 158) of each inner pocket 126 defined by the inner ring 154 andan aft portion (i.e., inner pocket aft segment 156) of each inner pocket126 defined by the inner band 102.

As illustrated in FIGS. 6A and 8, in some embodiments the inner band 102is truncated or shortened, and the inner ring 154 extends beyond aforward portion of the fairings 106. In the depicted embodiment, aflange 162 extends from the forward edge 160 of the inner band 102, anda flange 164 extends from an aft edge 166 of the inner ring 154. Eachflange 162, 164 defines a plurality of apertures 168, and when the innerring 154 is positioned against the forward edge 160 of the inner band102, apertures 168 of the inner band flange 162 align with apertures 168of the inner ring flange 164. A fastener 170, such as a bolt, grommet,rivet, or the like, may extend through the flanges 162, 164 within eachpair of aligned apertures 168 to attach or couple the inner ring 154 tothe inner band 102. As shown most clearly in FIG. 6A, the forward edge160 of the inner band 102 is defined axially aft or downstream of theleading edges 118 of the fairings 106, such that a joint defined wherethe inner ring 154 is positioned against the inner band 102 is aft ordownstream of the leading edge 118 of each fairing 106.

In other embodiments, as shown in FIG. 6B, the inner band 102 is nottruncated or shortened, and the inner ring 154 functions primarily toclose the forward end 130 of each inner pocket 126. In such embodiments,the inner ring 154 also provides additional structure to the inner band102, e.g., to provide support for the fairing assembly 100 at itsforward inner end. It will be appreciated that the inner ring 154 shownin FIG. 6B therefore may be similar to the outer ring 172 describedherein, for example, with respect to FIGS. 9-11.

Further, it will be appreciated that the outer band 104 and fairings 106of the embodiment of the fairing assembly 100 shown in FIGS. 6A-8 may bethe same as the outer band 104 and fairings shown in detail in FIGS. 4and 5. That is, the fairing assembly 100 illustrated in FIGS. 6A-8comprises an annular single piece outer band 104. The plurality offairings 106, an example of which is illustrated in FIG. 5, extend fromthe inner pockets 126 defined by the inner ring 154 and inner band 102to the outer pockets 128 defined by the outer band 104. Moreover, theinner band 102 and the inner ring 154 are generally conical in shape,like the inner band 102 shown in FIGS. 2A-4.

Like the embodiments shown in and described with respect to FIGS. 2A-5,the fairing assembly 100 of FIGS. 6A-8 may be pinned at the forward end110 of the assembly 100 to retain the assembly 100 in position withinthe gas turbine engine 10. As previously described, the radial pins 138,144 allow some radial movement, e.g., to compensate for different ratesof thermal growth in the radial direction R by components formed fromdifferent materials, while constraining the fairing assembly 100 axiallyand tangentially (or circumferentially). In some embodiments, however,the fairing assembly 100 may be bolted rather than pinned to, e.g., theturbine frame, as described above. In particular embodiments, theassembly 100 may be bolted to one or more flexible metal hangers, whichflex to compensate for the thermal growth of the metal hanger(s)relative to the assembly 100, which may be formed from a compositematerial and have a different rate of thermal expansion. In furtherembodiments, the fairing airfoils 106 may include a feature that allowsthe fairings 106 to be bolted to a hanger. Other means for retaining thefairing assembly 100 in its position within the engine 10 may be used aswell.

As described herein, the fairing assembly 100 illustrated in FIGS. 6A-8is compatible with a two-piece frame or similar frame designs. Moreparticularly, the fairing assembly 100 shown in FIGS. 6A-8 is suitablefor a turbine frame that allows the single piece annular outer band 104to be installed with respect to the frame, the outer ends 124 of thefairings 106 positioned in the outer pockets 128, and the single pieceannular inner band 102 and the inner ring 154 to be slid into positionsuch that the inner ends 122 of the fairings 106 are received in theinner pockets 126. Methods of assembling the fairing assembly 100 willbe described in greater detail below.

Referring now to FIGS. 9-13, another exemplary embodiment of the fairingassembly 100 will be described. FIG. 9 provides a schematiccross-section view of the fairing assembly 100, and FIG. 10 provides aperspective view of the fairing assembly 100. FIG. 11 provides anexploded perspective view of the outer band 104 and an outer ring 172 ofthe fairing assembly 100. FIG. 12 provides an exploded perspective viewof an axially split fairing 106, and FIG. 13 provides a schematiccross-section view of a circumferentially split fairing 106. Further,although not depicted in FIG. 9, it will be appreciated that the fairingassembly 100 will be installed with a turbine frame like the turbineframe shown and described with respect to FIGS. 2A and 2B, although theturbine frame of FIG. 9 may be configured differently from the turbineframe depicted in FIGS. 2A and 2B, as well as the turbine frame withwhich the fairing assembly 100 of FIGS. 6A-8 is used. For instance, thefairing assembly 100 depicted in FIGS. 9-13 may be used with a singlepiece or integral turbine frame, or similar frame design, rather than abolted turbine frame or the like as described above with respect toFIGS. 2A-5.

As depicted in FIGS. 9-11, the fairing assembly 100 may comprise anannular, single piece outer ring 172 in addition to the annular, singlepiece outer band 104. More specifically, as shown most clearly in FIG.11, in some embodiments of the outer band 104, each outer pocket 128 isopen at its forward end 134 and closed at its aft end 136. The outerband 104 includes a forward flange 174, and the outer ring 172 ispositioned radially inward of the forward flange 174, as illustrated inFIGS. 9 and 10. The outer ring 172 extends across the open forward ends134 of the outer pockets 128 to close the forward end 134 of each outerpocket 128; the outer band 104 is otherwise generally the same as theouter band 104 shown in FIGS. 2A-4. The outer ring 172 may be relativelythick along the radial direction R to provide structural support at theforward end 110 of the assembly 100; as shown in FIGS. 2A-4, 6, and 7,in other embodiments the outer band 104 is relatively thick at theforward end 110 to provide structural support to the assembly 100.Further, as illustrated in FIGS. 9 and 10, the same inner band 102 andinner ring 154 are used in the embodiment of fairing assembly 100 shownin FIGS. 9-12 as in the embodiment of fairing assembly 100 illustratedin FIGS. 6A, 7, and 8.

Moreover, like the embodiments shown in and described with respect toFIGS. 2A-5 and 6-8, the fairing assembly 100 of FIGS. 9-12 may be pinnedat the forward end 110 of the assembly 100 to retain the assembly 100 inposition within the gas turbine engine 10. A plurality of apertures 176may be defined in the forward flange 174 of the outer band 104 and aplurality of openings 178 may be defined in the outer ring 172. Theapertures 176 are defined in the outer band 104 and the openings 178 aredefined in the outer ring 172 such that when the outer ring 172 isassembled with the outer band 104, the outer band apertures 176 alignwith the outer ring openings 178. An outer pin 144 may be received ineach pair of aligned outer band apertures 176 and outer ring openings178 to retain the outer ring 172 with respect to the outer band 104.

Further, as previously described, the radial pins 138, 144 allow someradial movement, e.g., to compensate for different rates of thermalgrowth in the radial direction R by components formed from differentmaterials, while constraining the fairing assembly 100 axially andtangentially (or circumferentially). In some embodiments, however, thefairing assembly 100 may be bolted rather than pinned to, e.g., theturbine frame, as described above. In particular embodiments, theassembly 100 may be bolted to one or more flexible metal hangers, whichflex to compensate for the thermal growth of the metal hanger(s)relative to the assembly 100, which may be formed from a compositematerial and have a different rate of thermal expansion. In furtherembodiments, the fairing airfoils 106 may include a feature that allowsthe fairings 106 to be bolted to a hanger. Other means for retaining thefairing assembly 100 in its position within the engine 10 may be used aswell.

As described herein, the fairing assembly 100 illustrated in FIGS. 9-11is compatible with a single piece frame or similar frame designs. Moreparticularly, the fairing assembly 100 shown in FIGS. 9-11 utilizes twoannular pieces at each of the inner portion and outer portion of theassembly 100 such that the assembly 100 can be installed around thesingle piece frame. Further, referring particularly to FIG. 12, at leastone fairing 106 of the plurality of fairings 106 may be split axiallyinto a forward section 180 and an aft section 182 such that the fairing106 may be positioned around and encircle one of the struts 108. Thatis, because of the turbine frame design, the fairings 106 must be splitto be installed around the struts 108. The forward section 180 and aftsection 182 of each fairing 106 may abut along a first join line 184 anda second join line 186. More particularly, the forward section 180 andaft section 182 may each define a rabbet or one half of a half lapjoint. The forward and aft sections 180, 182 connect together along thefirst and second join lines 184, 186 to form the fairing 106. Further,it will be understood that the inner band 102 and outer band 104, inwhich the respective ends 122, 124 of the fairing 106 are received, helpkeep the forward and aft sections 180, 182 together (i.e., help keep theforward and aft sections 180, 182 from coming apart).

In other embodiments, as illustrated in FIG. 13, at least one fairing106 of the plurality of fairings 106 may be split circumferentially intoa first side section 188 and a second side section 190; one side section188, 190 may correspond to the pressure side and the other side section188, 190 may correspond to the suction side of the airfoil-shapedfairing 106. The first side section 188 and second side section of eachfairing 106 abut along a first joint 192 at the leading edge 118 and asecond joint 194 at the trailing edge 120. As shown in FIG. 13, each ofthe first joint 192 and second joint 194 are overlapping joints, e.g.,the second side section 190 overlaps the first side section 188 at thefirst joint 192 and the first side section 188 overlaps the second sidesection 190 at the second joint 194. More particularly, each of thefirst and second side sections 188, 190 define notches 196 along theleading 118, and the first and second side section 188, 190 fit togetheralong the notches 196 to form overlapping first joint 192. The firstside section 188 defines a notch 198 at the trailing edge 120, and thesecond side section 190 defines a projection 200 that fits within thenotch 198 to form overlapping second joint 194. Moreover, it will beappreciated that the inner band 102 and outer band 104, in which therespective ends 122, 124 of the fairing 106 are received, help keep thefirst and second side sections 188, 190 together (i.e., help keep thefirst and second side sections 188, 190 from coming apart).

As illustrated in FIGS. 5, 12, and 13, whether formed as a single piecestructure, split axially, or split circumferentially, each fairing 106defines a cavity 202, which may be sized to receive a strut 108 or otherstructural component. A fluid may be received in each fairing airfoilcavity 202 such that the internal pressure of the fairing 106 is higherthan the external pressure of the fairing 106. In the split fairingembodiments of FIGS. 12 and 13, the higher cavity pressure may help pushthe fairing sections 180, 182 and 188, 190 together for a tight sealbetween the sections. Further, as illustrated in FIG. 13, any leakagefrom the cavity 202 would be forced axially to the trailing edge 120 andsecond joint 194 of the fairing 106. Accordingly, a seal 204, such as aspline seal or the like, may extend along the second joint 194 betweenthe first side section 188 and second side section 190, e.g., betweenthe notch 198 and projection 200, to help reduce leakage from the higherpressure cavity 202. Other seals also may be included with eitherfairing 106 shown in FIGS. 12 and 13 to help reduce leakage from thecavity 202.

Turning now to FIG. 14, each inner pocket 126 and outer pocket 128 maybe built up to create a stop for each fairing 106. Further, each of theinner band 102 and outer band 104 may be built up at the aft end 112 ofthe fairing assembly 100 to provide an area for a seal at each of theinner band 102 and outer band 104. More specifically, each inner pocket126 comprises a lip 206 that extends about the inner pocket 126, and theinner end 122 of each fairing 106 is received within an inner pocket 126such that the inner end 122 contacts the lip 206. Similarly, each outerpocket 128 comprises a lip 208 that extends about the outer pocket 128,and the outer end 124 of each fairing 106 is received within an outerpocket 128 such that the outer end 124 contacts the lip 208. As such,the inner lip 206 acts as an inner stop for the fairing 106, and theouter lip 208 acts as an outer stop for the fairing 106, where the innerand outer lips 206, 208 help prevent radial and/or tangential (orcircumferential) slippage of the fairings 106. In some embodiments, aseal 210 such as a wire seal may extend around each of the inner pockets126 and outer pockets 128, e.g., in the inner and outer pockets 126, 128near where the inner and outer bands 102, 104 are built up to define thelips 206, 208 such that the seals 210 can contact each of the inner end122 and outer end 124 of the fairings 106 to form a seal between thefairings 106 and the bands 102, 104. Because the internal pressure ofthe fairings 106 is greater than the external or flow path pressure, theinternal pressure pushes the fairings 106 into the seals 210, whichhelps increase the effectiveness of the seals 210, i.e., creates a goodseal between the fairings 106 and the bands 102, 104. The split fairings106 shown in FIGS. 12 and 13 may have particularly good seals betweenthe fairings 106 and the bands 102, 104, as the higher internal pressureof the fairings 106 pushes the sections 180, 182 or 188, 190 into theseals 210 and bands 102, 104.

Further, an aft edge 212 of the inner band 102 may be built up, e.g.,may be thicker than the rest of the inner band 102, to define an innersurface 214. Likewise, an aft edge 216 of the outer band 104 may bebuilt up, e.g., may be thicker than the rest of the outer band 104, todefine an outer surface 218. Each of the inner surface 214 and outersurface 218 may provide a surface against which a seal, such as a pistonring seal or the like, may be positioned such that the fairing assembly100 is sealed at its aft end 112 along each of the inner band 102 andouter band 104.

FIGS. 15, 16, and 17 provide radial cross-section views of a portion ofa fairing end and a band pocket with a seal positioned therebetween,e.g., to provide sealing and wear protection between the fairing 106 andrespective band 102, 104, according to various exemplary embodiments ofthe present subject matter. Referring particularly to FIG. 15, a radialcross-section view is provided of a portion of an outer pocket 128 ofouter band 104 and outer end 124 of fairing 106 according to anexemplary embodiment of the present subject matter. As shown in FIG. 15,a seal 220 may be positioned between the outer end 124 of the fairing106 and the outer pocket 128 of the outer band 104. More particularly,the seal 220 includes a wear portion 222 that extends along a radialsurface 224 of the outer pocket 128. A retainer portion 226 extends fromthe wear portion 222 of the seal 220 into the outer band 104, e.g., toretain the seal 220 in position within the outer pocket 128. Theretainer portion 226 may extend about the entire seal 220, which mayextend about the entire perimeter of the outer pocket 128, or the seal220 may define a plurality of retainer portions 226 that are spacedapart from other another such that the retainer portions 226 arereceived within the outer pocket 128 at various locations about theouter pocket 128.

Further, the wear portion 222 of the seal 220 defines a planar wearsurface 228 against which the outer end 124 of the fairing 106 may bepositioned. That is, the wear surface 228 of the seal 220 is between theouter end 124 of the fairing 106 and the outer band 104 such that thefairing 106 may rub or slide against, or otherwise contact, the seal 220rather than the outer band 102, thereby helping prevent wear between thefairing 106 and outer band 104 in the area of the outer pocket 128.Additionally, the seal 220 includes a seal arm 230 that is compressedbetween the outer end 124 of the fairing 106 and the outer band 104.More specifically, the seal arm 230 extends between the lip 208 of theouter pocket 128 and a radially outermost surface 232 of the fairing106. The seal arm 230 is curved such that it has a generally serpentineor S-shaped cross-section. In the exemplary embodiment illustrated inFIG. 15, the seal arm 230 projects from the wear portion 222 in contactwith the outer pocket lip 208, then curves toward the surface 232 of thefairing 106 until the seal arm 230 contacts the surface 232. Thus, theseal 220 is configured to allow for radial thermal growth of the outerband 102, fairing 106, seal 220, and/or components surrounding and/orsupporting the fairing assembly 100, while also providing a seal andwear protection between the fairing 106 and outer pocket 128.

It will be appreciated that a plurality of seals 220 may be provided forthe fairing assembly 100. One seal 220 of the plurality of seals 220 mayextend within each outer pocket 128 such that the seal 220 is betweenthe outer pocket 128 and the outer end 124 of the fairing 106 receivedwithin the outer pocket 128. Further, seals 220 also may be used betweenthe inner end 122 of the fairing 106 and the inner pocket 126. In suchembodiments, the inner fairing end 122, inner pocket 126, and seal 220may be configured as shown in FIG. 15, with the seal arm 230 extendingradially outward rather than radially inward from the seal 220, i.e.,when used as an inner seal 220, the seal 220 will be flipped or rotatedabout the circumferential direction C from the view provided in FIG. 15.

Turning to FIG. 16, a radial cross-section view is provided of a portionof an outer pocket 128 of outer band 104 and outer end 124 of fairing106 according to another exemplary embodiment of the present subjectmatter. As shown in FIG. 16, a seal 234 may be positioned between theouter end 124 of the fairing 106 and the outer pocket 128 of the outerband 104. More particularly, the seal 234 defines a channel 236 intowhich the outer end 124 of the fairing 106 is received such that theseal 234 clips on or otherwise attaches to the fairing outer end 124. Inthe depicted embodiment, the fairing outer end 124 is notched along aninner surface 238 and an outer surface 240, and the seal 234 includesprojections 242 that are received in the notches 244 to attach the seal234 to the fairing outer end 124. Further, the seal 234 includes a wearportion 246 that defines a planar wear surface 248 that may bepositioned against the outer pocket 128. That is, the wear surface 248of the seal 234 is between the outer end 124 of the fairing 106 and theouter band 104 such that the seal 234 may rub or slide against, orotherwise contact, the outer band 104 rather than the fairing 106contacting the outer band 104, thereby helping prevent wear between thefairing 106 and outer band 104 in the area of the outer pocket 128.

Moreover, like the seal 220 shown in FIG. 15, the seal 234 includes aseal arm 250 that is compressed between the outer end 124 of the fairing106 and the outer band 104. More specifically, the seal arm 250 extendsbetween the radially outermost surface 232 of the fairing 106 and thelip 208 of the outer pocket 128. The seal arm 250 is curved such that ithas a generally C-shaped or U-shaped cross-section. In the exemplaryembodiment illustrated in FIG. 16, the seal arm 250 projects from theseal channel 236 and curves toward the outer pocket lip 208 until theseal arm 250 contacts the lip 208. As such, the seal 234 is configuredto allow for radial thermal growth of the outer band 102, fairing 106,seal 234, and/or components surrounding and/or supporting the fairingassembly 100, while also providing a seal and wear protection betweenthe fairing 106 and outer pocket 128.

It will be understood that a plurality of seals 234 may be provided forthe fairing assembly 100. One seal 234 of the plurality of seals 234 mayextend around the outer end 124 of each fairing 106 such that the seal234 is between the outer pocket 128 and the outer end 124 of the fairing106 received within the outer pocket 128. Further, seals 234 also may beused between the inner end 122 of the fairing 106 and the inner pocket126. In such embodiments, the inner fairing end 122, inner pocket 126,and seal 234 may be configured as shown in FIG. 16, with the seal arm250 extending radially inward toward the inner lip 206 rather thanradially outward toward the outer lip 208, i.e., when used as an innerseal 234, the seal 234 will be flipped or rotated about thecircumferential direction C from the view provided in FIG. 16.

Referring to FIG. 17, a radial cross-section view is provided of aportion of an outer pocket 128 of outer band 104 and outer end 124 offairing 106 illustrating another exemplary embodiment of a seal betweenthe fairing 106 and outer band 104. As shown in FIG. 17, a seal 252 maybe clipped on or otherwise attached to the outer lip 208 rather than thefairing outer end 124 as shown in FIG. 16 such that the seal 252 ispositioned between the fairing outer end 124 and the outer pocket 128 ofthe outer band 104. More particularly, the seal 252 defines a channel254 into which the lip 208 of the outer pocket 128 is received. An arm256 of the channel 254 extends along each of an inner surface 258 and anouter surface 260 of the lip 208. Further, the inner arm 256 of the seal252 functions as a wear portion of the seal 252, defining a planar wearsurface 262 that may be positioned against the outermost surface 232 ofthe fairing 106. That is, the wear surface 262 of the seal 252 isbetween the outer end 124 of the fairing 106 and the outer band 104 suchthat the fairing 106 may rub or slide against, or otherwise contact, theseal 252 rather than the outer band 104, thereby helping prevent wearbetween the fairing 106 and outer band 104 in the area of the outerpocket 128.

In addition, like the seals 220 and 234 shown in FIGS. 15 and 16, theseal 252 includes a seal arm 264 that is compressed between the outerend 124 of the fairing 106 and the outer band 104. More specifically,the seal arm 264 extends between the radial surface 224 of the outerpocket 128 and the sides 114, 116 of the fairing 106. The radial surface225 may define a recess 225 such that the seal arm 264 extends radiallywithin the recess 225. Further, the seal arm 264 is curved such that ithas a generally C-shaped or U-shaped cross-section. In the exemplaryembodiment illustrated in FIG. 17, the seal arm 264 projects from theinner channel arm 256 along the radial surface 224 and curves toward theouter end 124 of the fairing until the seal arm 264 contacts the side114, 116 of the fairing 106; it will be appreciated that the seal 252and seal arm 264 extend about the fairing outer end 124 such that oneportion of the seal arm 264 contacts the pressure side 114 and theremaining portion contacts the suction side 116. Accordingly, the seal234 is configured to allow for axial and circumferential thermal growthof the outer band 102, fairing 106, seal 252, and/or componentssurrounding and/or supporting the fairing assembly 100, while alsoproviding a seal and wear protection between the fairing 106 and outerpocket 128.

It will be understood that a plurality of seals 252 may be provided forthe fairing assembly 100. One seal 252 of the plurality of seals 252 mayextend around the lip 208 of each outer pocket 128 such that the seal252 is between the outer pocket 128 and the outer end 124 of the fairing106 received within the outer pocket 128. Further, seals 252 also may beused between the inner end 122 of the fairing 106 and the inner pocket126. In such embodiments, the inner fairing end 122, inner pocket 126,and seal 252 may be configured substantially as shown in FIG. 17.

Turning now to FIGS. 18A, 18B, and 19, features for pinning the fairingairfoils 106 in the fairing assembly 100 will be described. Asillustrated in FIGS. 18A and 18B, a pin 266 having a retention member268 may be inserted into an aperture 270 in each fairing 106. Moreparticularly, a first end 270 of a pin 266 may be inserted into anaperture 272 on each side 114, 116 of each fairing 106. In embodimentsin which the fairing 106 is split into forward and aft sections 180,182, the pins 266 may help hold the sections 180, 182 together, asillustrated in FIG. 18B. Referring to FIG. 19, the apertures 272 may bedefined in the inner end 122 of each fairing 106 such that the pins 266pin the fairings 106 to the inner band 102. The inner band 102 defines agroove 274 in which a second end 276 of each pin 266 is received. Asshown in FIG. 19, the pinned fairings 106 may be best suited for usewith the fairing assemblies 100 utilizing an inner band 102 and innerring 154, e.g., as described with respect to FIGS. 6A-10. In suchembodiments, the inner band 102 may be positioned within the engine 10,then the fairings 106 having pins 266 inserted into apertures 272therein may be installed such that the fairing inner ends 122 arereceived in the inner pocket aft segments 156 and the pin second ends276 are received in the groove 274. The inner ring 154 may then bemaneuvered into position to enclose the fairing inner ends 122 in innerpockets 126 (the forward segments 158 of which are defined by the innerring 154) and close the groove 274 such that the pin second ends 276 areretained in the groove 274. The retention member 268 helps retain thepins 266 in the apertures 272 and minimize circumferential movement ofthe pins 266.

In alternative embodiments, the groove 274 may be defined in the innerring 154 rather than the inner band 102 (with the assembly methodaltered as needed to properly assemble such configuration) such that thepin second ends 276 are received in the inner ring 154 rather than theinner band 102. In still other embodiments, the groove 274 may bedefined in part by the inner band 102 and in part by the inner ring 154such that the pin second ends 276 are secured between the inner band 102and inner ring 154. Even further, in some embodiments the fairings 106may be pinned at the outer ends 124 rather than the inner ends 122, withthe apertures 272 being defined in the fairing outer ends 124 and thegroove 274 (or, as appropriate, apertures for receipt of the pin secondends 276) being defined in the outer band 104.

As further illustrated in FIG. 19, a seal may extend around each of theinner end 122 and outer end 124 of the fairing 106. In some embodiments,the inner seal 278 may be a wire seal or the like having a roundcross-sectional shape, and the outer seal 280 may be a piston seal orthe like having a square cross-sectional shape. The square cross-sectionseal 280 may be suitable for planar geometry, which may exist at theouter ends 124, while the round cross-section seal 278 may be suitablefor a more complex geometry, which may exist at the inner ends 122. Ofcourse, any suitable seal may be used, including the exemplary seals220, 234, 252 described with respect to FIGS. 15-17.

FIGS. 20A, 20B, and 21-24 provide schematic cross-section views ofvarious grommet and fastener configurations according to exemplaryembodiments of the present subject matter. The grommet and fastenerconfigurations, e.g., may be used to secure portions of the fairingassembly 100 to one another, within the engine 10, etc. Referring toFIGS. 20A and 20B, a grommet configuration is provided, which may beused in a blind or non-through hole, e.g., to line the hole and therebyprotect the component in which the hole is defined. More particularly, anon-through hole 282 may be defined in a component, e.g., the forwardportion 148 of the outer band 104 as shown in FIGS. 2A and 2B, and thehole 282 may be machined such that it tapers outward along its side 284near its closed distal end 286. That is, as shown in FIG. 20A, thecross-section of the hole 282 is larger near the closed distal end 286than near its open proximal end 288. A grommet 290 having an angled end292 is inserted into the hole 282 with the angled end 292 anglinginward, i.e., toward an axial centerline 294 of the grommet 290. A tool296 may be inserted into the grommet 290 to push the angled end 292outward as shown in FIG. 20B, such that the angled end 292 flares orangles outward after insertion of the tool 296 and fits within thetapered portion, i.e., distal end 286, of the hole 282. It will beappreciated that, in other embodiments, the angled grommet 290 may beused to hold two components together rather than merely lining the hole282, e.g., to protect the component defining the hole 282. Further,rather than using a tool 296 to push or flare out the angled end 292 ofthe grommet 290, a fastener such as a screw, bolt, pin, or the like maybe inserted into the grommet 290 to push or flare out the grommet angledend 292.

Turning to FIGS. 21 and 22, in some embodiments, multiple grommets and apin may be used to secure two components to one another. As shown inFIG. 21, a grommet 298 may be inserted and secured into the non-throughhole 282 in a first component 300 and a swaged grommet 302 may beattached in an opening 304 of a second component 306; the secondcomponent 306 is being coupled to the first component 300. A pin 308having a head 310 is inserted into the opening 312 defined by thegrommets 298, 302, and the pin 308 may be welded to the swaged grommet302. As shown in FIG. 22, rather than inserting the grommet 298 in anon-through hole such as hole 282, the grommet 298 may be secured in athrough hole 314. Further, a headless pin 316 may be used to couple thefirst and second components 300, 306, and the grommet 298 may include abottom portion 318 to prevent the pin 316 from slipping through theopening 312 defined by the grommets 298, 302. Further, the pin 316 maybe welded to the grommet 302 to secure the pin 316 in place.

Referring now to FIG. 23, a flared insert or grommet 320 may be used inthe first component 300. More specifically, the grommet 320 is flaredoutward at its distal end 322, such that the distal end 322 has a largercross-section than a body 324 of the grommet 320, and the hole 282 or314 in which the grommet 320 is received is likewise flared or angled toreceive the grommet distal end 322. The grommet 320 may include ananti-rotation feature 326, such as a collar or the like that seats inthe first component 300, at a proximal end 328 of the grommet 320 toprevent the grommet 320 from rotating within the hole 282 or 314 inwhich the grommet 320 is secured. A fastener 330, such as a pin or bolt,may be inserted through the second component 306 and into the grommet320 to secure the first and second components 300, 306 together.Further, a washer 332 may be used, e.g., to prevent wear between thefastener 330 and the second component 306.

FIG. 24 illustrates another embodiment of a grommet and fastenerconfiguration for securing two components to one another. In theembodiment of FIG. 24, the grommet 320 extends within both the firstcomponent 300 and the second component 306. As shown in FIG. 23, thegrommet 320 includes a flared distal end 322 that is received within aflared hole 282 or 314 in the first component 300. The body 324 of thegrommet 320 extends through the second component 306, and the proximalend 328 may be swaged around the fastener 330 to retain the fastener inthe grommet 320. In other embodiments, the fastener 330 may be welded tothe grommet 320, in addition to or as an alternative to swaging thegrommet proximal end 328.

FIG. 25 provides a flow diagram illustrating a method 2500 forassembling a fairing assembly 100 in a gas turbine engine, such asturbofan engine 10, according to an exemplary embodiment of the presentsubject matter. The method 2500 includes installing an annular, singlepiece inner band 102 in the gas turbine engine, as shown at 2510. Theinner band 102 defines a plurality of inner pockets 126 as describedherein, and the method 2500 includes, as shown at 2520, inserting aninner end 122 of each of a plurality of fairings 106 into an innerpocket 126 of the plurality of inner pockets 126. Next, as illustratedat 2530, the method 2500 includes sliding an annular, single piece outerband 104 into position with respect to the plurality of fairings 106such than an outer end 124 of each of the plurality of fairings 106 isreceived in an outer pocket 128 of a plurality of outer pockets 128defined by the outer band 104. In exemplary embodiments, the outer band104 slides into position from a forward end 110 of the fairing assembly100 toward an aft end 112 of the assembly 100. Optionally, method 2500also may include securing the fairings 106 to one or both of the innerband 102 and outer band 104 via pins, bolts, brazing, bonding, or anyother suitable attachment means. For example, the fairings 106 may besecured to the inner band 102 and/or outer band 104 once the outer band104 is in position with respect to the fairings 106.

Method 2500 may be used to assemble the fairing assembly 100 asdescribed with respect to FIGS. 2A-5 with a turbine frame having abolted frame design, although method 2500 may be used with other framedesigns as well. In embodiments in which the fairing assembly is usedwith a bolted turbine frame, the method 2500 may further includeinstalling the struts 108 through the inner band 102, outer band 104,and fairings 106 of the fairing assembly 100, as shown at 2540, althougha strut 108 need not extend through every fairing 106. Next, asillustrated at 2550 and 2560, the struts 108 are secured to the innerhub 140 and the outer case 146, e.g., by bolting the struts 108 to thehub 140 and case 146. As depicted in FIG. 2A, the fairing assembly 100may be pinned at the outer side of its forward end 110, and thus, themethod 2500 includes at 2570 inserting at least one pin 144 through theouter case 146 and into an opening 152 in the outer band 104. Theopening 152 may be a blind hole as described herein. The method 2500also may include steps for securing the pin 144 with respect to thefairing assembly 100. Further, as illustrated in FIG. 2B, the fairingassembly 100 also may be pinned at the inner side of its forward end110, such that the method 2500 includes at 2570 inserting at least onepin 138 through the inner hub 140 and into an opening 150 in the innerband 102 to pin the fairing assembly 100 on its inner side as well asits outer side.

Other exemplary assembly methods also are provided. FIG. 26 provides aflow diagram illustrating a method 2600 for assembling a fairingassembly 100 in a gas turbine engine, such as turbofan engine 10,according to another exemplary embodiment of the present subject matter.The method 2600 includes, as shown at 2610, installing one fairing 106of a plurality of fairings 106 over each strut 108 of the turbine frame.Then, the method 2600 comprises installing an annular, single pieceouter band 104 in the gas turbine engine, as shown at 2620. The outerband 104 defines a plurality of outer pockets 128 as described herein,and the method 2600 includes, as shown at 2630, inserting an outer end124 of each of a plurality of fairings 106 into an outer pocket 128 ofthe plurality of outer pockets 128. Next, as illustrated at 2640, themethod 2600 includes sliding an annular, single piece inner band 102into position with respect to the plurality of fairings 106 such thatthe inner end 122 of each of the plurality of fairings 106 is receivedin an inner pocket aft segment 156 of a plurality of inner pocket aftsegments 156 defined by the inner band 102.

Then, as shown at 2650, the method 2600 comprises positioning an innerring 154 at a forward edge 160 of the inner band 102. As describedherein, the two pieces, inner band 102 and inner ring 154, of thefairing assembly 100 allow the fairing assembly 100 to be installedaround the turbine frame. The inner pocket aft segments 156 of the innerband 102 form open inner pockets 126, i.e., pockets that are open at aforward end 130 and closed at an aft end 132, and the inner ring 154 ispositioned at the forward edge 160 of the inner band 102 to close theforward ends 130. As described herein, the inner ring 154 defines aplurality of inner pocket forward segments 158, and together the innerpocket forward segments 158 and inner pocket aft segments 156 define theinner pockets 126 that encircle the inner ends 122 of the plurality offairings 106. The method 2600 further may include, as shown at 2660,joining the inner ring 154 to the inner band 102, e.g., fastening aninner ring flange 164 to an inner band flange 162 with suitablefasteners 170. Optionally, method 2600 also may include securing thefairings 106 to one or both of the inner band 102 and outer band 104 viapins, bolts, brazing, bonding, or any other suitable attachment means.For example, the fairings 106 may be secured to the inner band 102and/or outer band 104 once the inner band 102 is in position withrespect to the fairings 106. Method 2600 may be used to assemble thefairing assembly 100 as described with respect to FIGS. 6A-8 with aturbine frame having a two-piece frame design, but method 2600 may beused with other frame designs as well.

FIG. 27 provides a flow diagram illustrating a method 2700 forassembling a fairing assembly 100 in a gas turbine engine, such asturbofan engine 10, according to yet another exemplary embodiment of thepresent subject matter. Method 2700 may be utilized to assemble thefairing assembly 100 as described with respect to FIGS. 9-13 with aturbine frame having a single or one-piece frame design, although method2700 also may be used with other frame designs. As such, the method 2700includes installing the inner band 102, outer band 104, and fairings 106around the turbine frame. Accordingly, the open inner and outer pockets126, 128 configuration may be utilized, where the inner ring 154 andouter ring 172 are provided to close the inner and outer pockets 126,128.

Referring to FIG. 27, as shown at 2710, the method 2700 includesinstalling an annular, single piece outer band 104 in the gas turbineengine. The outer band 104 defines a plurality of outer pockets 128 asdescribed herein. Next, as illustrated at 2720, the method 2700 includespositioning the outer ring 172 at the forward flange 174 of the outerband 104, as shown at 2720, and joining the outer ring 172 to the outerband 104, as depicted at 2730. Thus, the two pieces, outer band 104 andouter ring 172, of the outer portion of the fairing assembly 100 areinstalled around the turbine frame.

Then, as illustrated at 2740, a plurality of first fairing sections maybe assembled with a plurality of second fairing sections to form aplurality of fairings 106. As described with respect to FIGS. 12 and 13,the first and second fairing sections may be forward and aft sections180, 182 or first side and second side sections 188, 190 of a splitfairing 106. The fairings 106 are split such that they may be installedaround struts 108 and/or other components of the single piece turbineframe. After the first and section fairing sections are assembled toform the fairings 106, each fairing 106 is slid into an outer pocket 128defined in the outer band 104, as shown at 2750.

The method 2700 also includes sliding an annular, single piece innerband 102 into position with respect to the assembled fairings 106, asshown at 2760. The inner band 102 defines a plurality of inner pockets126. Next, the method 2700 includes positioning the inner ring 154 atthe forward edge 160 of the inner band 102, as shown at 2770, andjoining or fastening the inner ring 154 to the inner band 102, as shownat 2780. Therefore, the two pieces, inner band 102 and inner ring 154,of the inner portion of the fairing assembly 100 are installed aroundthe turbine frame. As described herein, the inner ring 154 closes theinner pockets 126 and the outer ring 172 closes the outer pockets 128.The inner ring 154 may be fastened to the inner band 102 using anyappropriate fastener, and the outer ring 172 may be joined to the outerband 104, e.g., by pinning the outer ring 172 to the outer band 104.Optionally, method 2700 also may include securing the fairings 106 toone or both of the inner band 102 and outer band 104 via pins, bolts,brazing, bonding, or any other suitable attachment means. For example,the fairings 106 may be secured to the inner band 102 and/or outer band104 once the inner band 102 is in position with respect to the fairings106.

Further, as described herein, the inner band 102, outer band 104, andfairings 106 may be formed from a CMC material. It will be appreciatedthat the inner ring 154 and outer ring 172 also may be formed from a CMCor other suitable composite and that the split fairings 106 may beformed from a CMC or other composite. However, the inner band 102, outerband 104, fairings 106, inner ring 154, and/or outer ring 172 also maybe formed from any suitable material. Specific processing techniques andparameters for forming the components of the fairing assembly 100 willdepend on the particular composition of the materials. For example,silicon CMC components may be formed from fibrous material that isinfiltrated with molten silicon, e.g., through a process typicallyreferred to as the Silcomp process. Another technique of manufacturingCMC components is the method known as the slurry cast melt infiltration(MI) process. Other techniques for forming CMC components includepolymer infiltration and pyrolysis (PIP) and oxide/oxide processes.Components may also be fabricated from a carbon fiber reinforced siliconcarbide matrix (C/SiC) CMC, which is processed using chemical vaporinfiltration.

Still further, in some embodiments, one or more components of thefairing assembly 100 may be formed using a suitable additivemanufacturing technique or process, such as Fused Deposition Modeling(FDM), Selective Laser Sintering (SLS), 3D printing such as by inkjetsand laserjets, Sterolithography (SLA), Direct Selective Laser Sintering(DSLS), Electron Beam Sintering (EBS), Electron Beam Melting (EBM),Laser Engineered Net Shaping (LENS), Laser Net Shape Manufacturing(LNSM), Direct Metal Deposition (DMD), Digital Light Processing (DLP),Direct Selective Laser Melting (DSLM), Selective Laser Melting (SLM),Direct Metal Laser Melting (DMLM), and other known processes. Inaddition to using a direct metal laser sintering (DMLS) or direct metallaser melting (DMLM) process where an energy source is used toselectively sinter or melt portions of a layer of powder, it should beappreciated that according to alternative embodiments, the additivemanufacturing process may be a “binder jetting” process. In this regard,binder jetting involves successively depositing layers of additivepowder in a similar manner to a DMLS or DMLM process. However, insteadof using an energy source to generate an energy beam to selectively meltor fuse the additive powders, binder jetting involves selectivelydepositing a liquid binding agent onto each layer of powder. The liquidbinding agent may be, for example, a photo-curable polymer or anotherliquid bonding agent. Other suitable additive manufacturing methods andvariants are intended to be within the scope of the present subjectmatter.

The additive manufacturing processes described herein may be used forforming components using any suitable material. For example, thematerial may be plastic, metal, concrete, ceramic, polymer, epoxy,photopolymer resin, or any other suitable material that may be in solid,liquid, powder, sheet material, wire, or any other suitable form. Morespecifically, according to exemplary embodiments of the present subjectmatter, the additively manufactured components described herein may beformed in part, in whole, or in some combination of materials includingbut not limited to pure metals, nickel alloys, chrome alloys, titanium,titanium alloys, magnesium, magnesium alloys, aluminum, aluminum alloys,iron, iron alloys, stainless steel, and nickel or cobalt basedsuperalloys (e.g., those available under the name Inconel® availablefrom Special Metals Corporation). These materials are examples ofmaterials suitable for use in the additive manufacturing processesdescribed herein, and may be generally referred to as “additivematerials.”

Moreover, the additive manufacturing processes disclosed herein allows asingle component to be formed from multiple materials. Thus, thecomponents described herein may be formed from any suitable mixtures ofthe above materials. For example, a component may include multiplelayers, segments, or parts that are formed using different materials,processes, and/or on different additive manufacturing machines. In thismanner, components may be constructed that have different materials andmaterial properties for meeting the demands of any particularapplication. Further, although the components described herein may beconstructed entirely by additive manufacturing processes, it should beappreciated that in alternate embodiments, all or a portion of thesecomponents may be formed via casting, machining, a CMC component processas described herein, and/or any other suitable manufacturing process.Indeed, any suitable combination of materials and manufacturing methodsmay be used to form these components.

Accordingly, the embodiments described above provide a variety ofbenefits. For instance, the embodiments of the fairing assemblydescribed herein utilize single piece inner and outer bands, whichreduce leakage and pressure losses, as well as part count, manufacturingcomplexity, and manufacturing cost. Further, the split fairing assemblyas described herein is adapted for use with a variety of turbine frameconfigurations, including a single piece frame. Enabling use of thesingle piece turbine frame allows a reduced frame weight and reducedframe cost. Moreover, the fairing assembly embodiments described hereinmay be formed from a CMC material, which has a reduced weight andincreased temperature capability compared with other fairing assemblies.In addition, separating the fairing airfoils from the bands allows forrelative thermal expansion, which reduces thermal stress in the fairingassembly and allows the design to be more defect tolerant. Thus, theabove described embodiments result in commercial advantages such asreduced frame aerodynamic losses and manufacturing costs and allow forincreased operating temperatures and efficiency. Other advantages of thesubject matter described herein also may be realized by those ofordinary skill in the art.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. A fairing assembly for a gas turbine engine,comprising: a plurality of fairings, each fairing having an inner endradially spaced apart from an outer end, each fairing extending axiallyfrom a leading edge to a trailing edge; an annular inner band defining aplurality of inner pockets, each inner pocket shaped complementary tothe inner end of each fairing, each inner pocket having a forward endand an aft end; and an annular outer band defining a plurality of outerpockets, each outer pocket shaped complementary to the outer end of eachfairing, each outer pocket having a forward end and an aft end, whereinthe inner band is a single piece structure, wherein the outer band is asingle piece structure, and wherein the inner end of each fairing isreceived with an inner pocket of the plurality of inner pockets and theouter end of each fairing is received within an outer pocket of theplurality of outer pockets.
 2. The fairing assembly of claim 1, whereinthe inner band surrounds each inner pocket such that each inner pocketis closed at the forward end and the aft end.
 3. The fairing assembly ofclaim 1, wherein the outer band surrounds each outer pocket such thateach outer pocket is closed at the forward end and the aft end.
 4. Thefairing assembly of claim 1, wherein each inner pocket is open at theforward end and closed at the aft end.
 5. The fairing assembly of claim4, further comprising: an inner ring positioned against a forward edgeof the inner band to close the forward end of each inner pocket.
 6. Thefairing assembly of claim 4, wherein each outer pocket is open at theforward end and closed at the aft end.
 7. The fairing assembly of claim6, further comprising: an outer ring positioned at a forward flange ofthe outer band to close the forward end of each outer pocket.
 8. Thefairing assembly of claim 1, wherein the plurality of fairings, theinner band, and the outer band are each formed from a ceramic matrixcomposite material.
 9. A fairing assembly for a gas turbine engine,comprising: a plurality of fairings, each fairing having an inner endradially spaced apart from an outer end, each fairing extending axiallyfrom a leading edge to a trailing edge; an inner ring defining aplurality of inner pocket forward segments; an annular inner banddefining a plurality of inner pocket aft segments, the inner ringpositioned against a forward edge of the inner band such that the innerpocket forward segments and inner pocket aft segments form a pluralityof inner pockets, each inner pocket shaped complementary to the innerend of each fairing, each inner pocket having a forward end and an aftend; and an annular outer band defining a plurality of outer pockets,each outer pocket shaped complementary to the outer end of each fairing,each outer pocket having a forward end and an aft end, wherein the innerring is a single piece structure, wherein the inner band is a singlepiece structure, wherein the outer band is a single piece structure, andwherein the inner end of each fairing is received with an inner pocketof the plurality of inner pockets and the outer end of each fairing isreceived within an outer pocket of the plurality of outer pockets. 10.The fairing assembly of claim 9, further comprising: an outer ring,wherein the outer ring is a single piece structure, wherein each outerpocket is open at the forward end and closed at the aft end, and whereinthe outer ring is positioned at a forward flange of the outer band toclose the forward end of each outer pocket.
 11. The fairing assembly ofclaim 9, wherein at least one fairing is split axially and comprises aforward section and an aft section.
 12. The fairing assembly of claim 9,wherein at least one fairing is split circumferentially and comprises afirst side section and a second side section.
 13. The fairing assemblyof claim 9, wherein each inner pocket comprises a lip that extends aboutthe inner pocket, and wherein the inner end of each fairing is receivedwithin an inner pocket such that the inner end contacts the lip.
 14. Thefairing assembly of claim 9, wherein each outer pocket comprises a lipthat extends about the outer pocket, and wherein the outer end of eachfairing is received within an outer pocket such that the outer endcontacts the lip.
 15. The fairing assembly of claim 9, furthercomprising: a plurality of seals, wherein a seal of the plurality ofseals is positioned between the outer end of each fairing and the outerband, wherein each seal of the plurality of seals comprises a curved armthat is compressed between the outer end of the fairing and the outerband, and wherein each seal of the plurality of seals comprises at leastone planar wear surface between the outer end of the fairing and theouter band.
 16. The fairing assembly of claim 9, further comprising: aplurality of seals, wherein a seal of the plurality of seals ispositioned between the inner end of each fairing and the inner band,wherein each seal of the plurality of seals comprises a curved arm thatis compressed between the inner end of the fairing and the inner band,and wherein each seal of the plurality of seals comprises at least oneplanar wear surface between the inner end of the fairing and the innerband.
 17. The fairing assembly of claim 9, wherein the plurality offairings are pinned at the inner end of each fairing with a plurality ofpins, each pin of the plurality of pins extending between the inner endof one fairing of the plurality of fairings and the inner band.
 18. Thefairing assembly of claim 9, wherein the plurality of fairings, theinner band, and the outer band are each formed from a ceramic matrixcomposite material.
 19. The fairing assembly of claim 9, furthercomprising: a non-through hole defined in a component of the fairingassembly, the non-through hole tapering outward along its side near itsclosed distal end; and an angled grommet inserted in the non-throughhole, the angled grommet having an angled end fitting within the tapereddistal end of the non-through hole.
 20. The fairing assembly of claim 9,further comprising: a non-through hole defined in a first component; anopening defined in a second component; a grommet secured in thenon-through hole; a swaged grommet secured in the opening; and a pininserted into an opening defined by the grommet and the swaged grommet,wherein the pin in welded to the swaged grommet.
 21. The fairingassembly of claim 20, wherein the grommet is a flared grommet that has adistal end with a larger cross-section than a body of the grommet. 22.The fairing assembly of claim 9, further comprising: a non-through holedefined in a first component; an opening defined in a second component;and a grommet secured in the non-through hole and the opening, whereinthe grommet is a flared grommet that has a distal end with a largercross-section than a body of the grommet, wherein the body of thegrommet extends through the opening in the second component, and whereina proximal end of the grommet is swaged around a fastener inserted inthe grommet to retain the fastener with respect to the grommet.
 23. Amethod for assembling a fairing assembly in a gas turbine engine, themethod comprising: installing an annular inner band in the gas turbineengine, the inner band defining a plurality of inner pockets; insertingan inner end of each of a plurality of fairings into an inner pocket ofthe plurality of inner pockets; and sliding an annular outer band intoposition with respect to the plurality of fairings such than an outerend of each of the plurality of fairings is received in an outer pocketof a plurality of outer pockets defined by the outer band, wherein theinner band is a single piece structure, and wherein the outer band is asingle piece structure.
 24. The method of claim 23, wherein the outerband slides into position from a forward end of the fairing assemblytoward an aft end of the fairing assembly.